In axial flow gas turbines, the rotating compressor comprises one or more bladed discs (each constituting a "stage") mounted on a shaft which is supported at spaced points within the compressor housing or shroud. The turbines are assembled with a clearance gap between the rotor elements and the surrounding shroud to allow for differential thermal expansion between the various elements and/or minor displacement of the axis of rotation of the shaft due to operating loads.
However, to minimize losses of efficiency due to recirculation, the clearance gap should be designed to be as small as possible during operation.
This is especially true for small, highly-loaded, low-aspect-ratio turbines which are extremely sensitive to tip leakage losses. The clearance between the blade tip and the shroud is relatively greater when the blade length is small and thus has more of an effect on turbine performance. This presents a problem because of the temperature rises in the device, the differences in the coeffiecient of expansion of the various parts causes the gap size to change. It is thus necessary either to leave a large enough gap to allow for the expansion at all extremes of operating temperature or to provide for temporary or limited rubbing of the rotating and stationary parts during certain transient conditions while providing some means for preventing damage to the parts.
The prior art has tried several different approaches to solving this problem. One approach is to try to maintain both the shroud and rotor components at nearly the same temperature so that they will expand at the same rate and thereby maintain a constant running clearance. See, for example, U.S. Pat. No. 3,039,737.
Another approach is to provide an aerodynamic seal between the shroud and blades by extending the blade tip into a circumferential trench formed into the shroud and/or by attaching devices to the blade tips to direct gases away from the clearance gap. See, for example, U.S. Pat. Nos. 2,927,724, 3,583,824 and 3,701,536.
Yet another approach is allow the shroud, or a portion thereof, to be deformed, in a non-destructive manner, by the rotating components themselves so that only enough clearance is formed to accommodate the thermal expansion experiences in a particular engine. This latter approach was investigated further during development of the present invention.
One method which allows the shroud to be deformed into close running relationship with the rotating blades involves providing a fragile metallic honeycomb or cellular structure on the interior of the shroud and allowing the rotation of the blades cut a close fitting path through the fragile structure. See, for example, U.S. Pat. Nos. 3,689,971, 4,063,742, 4,526,509 and 4,652,209.
Another method involves coating the shroud interior with a soft or porous metal layer so that, again, the rotating blades can cut, or abrade, a path through the material. See, for example, U.S. Pat. Nos. 4,664,973 and 4,671,735.
U.S. Pat. No. 2,742,224 to F. M. Burhans for a "Compressor Casing Lining" appears to teach a shroud coating material which has a very sharp, but low, melting temperature so that any frictional heat due to rubbing will cause the coating to immediately melt and pass through the turbine without damage. Suggested materials are: indium, tin, cadmium, lead, zinc, and certain aluminum alloys.
U.S. Pat. No. 3,836,156 to H. B. Dunthorne for an "Ablative Seal" appears to teach a very similar concept except that the materials suggested are brazing alloys useful at higher temperatures.
U.S. Pat. Nos. 4,405,284, 4,460,311 and 4,669,955 teach the use of hard, brittle ceramic materials, including a honeycomb structure, which may be abraded or worn away during the initial run-in of a new turbine assembly. Suggested compositions include zirconia, ZrO.sub.2, MgO, and alumina or the like.
Several problems still exist in providing an effective and inexpensive abradable material. For example, the porous metals are difficult to attach securely to the base material of the shroud and are also often degraded by the absrasive and/or erosive action of the hot gas stream. The honeycomb material must be very fragile so as not to damage the blades but yet it must be substantial enough to be handled during manufacture and installation without deforming.
Most importantly, both types of seals suffer the limitation that the very high local temperature generated at the line of contact may be sufficient to flow the surface material so that when the rubbing ceases, a hard skin is formed which could damage the rotating blades at the next contact.
Thus, there is a need in this art for an improved abradable sealing material and structure for allowing its use in gas turbines.
None of the foregoing suggest the structure and composition of the present invention. Generally similar ceramic compositions are, of course, well known for other uses. See, for example, U.S. Pat. No. 4,252,408.